This application relates to segmented cooling cavities for use within a turbine component having an air cooled blade outer air seal.
Gas turbine engines are known, and typically include a compressor compressing air and delivering it downstream into a combustion section. The air is mixed with fuel in the combustion section and combusted. Products of this combustion pass downstream over turbine rotors, driving the turbine rotors to rotate and create power.
Typically, the turbine rotors include a plurality of removable blades. Blade outer air seals are positioned adjacent a radially outer portion of the blades, to provide a tight clearance across the rotors such that air is restricted to flow over the rotors, rather than bypassing them. These blade outer air seals are subjected to a very harsh and hot environment due to the products of combustion. Thus, it is known to deliver cooling air to blade outer air seals.
A blade outer air seal typically has a leading edge and a trailing edge, defined in terms of the turbine gas flow direction. The pressure across the blade outer air seal drops since the products of combustion are transferring their pressure into energy to drive the turbine rotors. Thus, at the leading edge, the pressure is significantly higher than at the trailing edge. As an example, the pressure could be as much as twice as high as the leading edge as it is at the trailing edge.
This causes some issues with regard to the flow of cooling air within the blade outer air seal. In some prior art blade outer air seals, a single large cooling channel extends across the entire axial length of the seal, from leading to trailing edge. In such an arrangement, less air will exit at the leading edge compared to the trailing edge. This is because the cooling air will seek the lowest pressure, and will thus tend to flow more toward the trailing edge.
One solution to this problem has been the use of a plurality of separate cooling channels spaced along the length of the airfoil. These separate channels do limit the effect of the pressure differential between leading edge and the trailing edge. However, the use of the separate channels complicates the manufacture of the airfoil.
Most components for gas turbine engines containing an airfoil are formed by a loss core molding process. In such a process, a core is initially formed of a particular material. That core is inserted into a mold, and molten metal is directed into the mold and around the core. After the molten metal has hardened, the material of the core is leached away, leaving a cavity where the core once sat. This is typically the manner in which cooling channels are formed. To form the separate channels as mentioned above, the core must have a plurality of separate core members, or alternatively, a plurality of spaced core fingers. The use of several cores is complex, and the use of a single core with separate spaced fingers is not as structurally robust as may be desirable.